The present invention relates to the structure of a wall bounding the combustion chamber of a gas turbine engine, more particularly such a structure having a double wall construction.
Military and civilian use turbojet engines have used ever increasing compression ratios in their compressors which generate higher temperature gases at the high pressure compressor output, the combustion chamber and the high pressure turbine. Accordingly, the combustion chambers of these engines must be appropriately cooled because, as their output increases, the air flow available for cooling decreases.
Present gas turbine engine combustion chambers may be comprised of a double wall construction using internal tiles to minimize heat transfer from the combustion gases to the combustion chamber wall. Such tiles may be made of a ceramic material, such as SiC/SiC. Because such materials have little thermal conductivity, high cooling is required. It is furthermore known that the temperatures near the combustion chamber exit are critical for maximum engine performance. Thus, effective cooling of the combustion chamber while lowering the air flow necessary for such cooling is imperative.